Application of carbon fiber mesh for space and airborne platform applications

ABSTRACT

A solar array including a carbon fiber mesh substrate is provided. The solar panels including a series of hybrid structures formed along the common and continuous substrate by sandwiching the substrate between a series of discontinuous upper and lower support layers. In order to construct the solar panels having such hybrid structure, a series of top support layers or upper face sheets is disposed on a upper surface of the substrate and between the folding sections. The solar cells are placed on top of the upper face sheets. Similarly, a series of lower support layers or lower face sheets are disposed on a bottom surface of the substrate and between the folding sections. The folding sections are the regions where the upper and lower face sheets discontinue and expose the underlying substrate. The solar array can be folded along the folding section when a bending force is applied over one of the hybrid structures.

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This application claims the benefit of and is a continuation ofU.S. patent application bearing Ser. No. 09/654,924 filed Sep. 5, 2000.

BACKGROUND OF THE INVENTION

[0002] 1. Field of the Invention

[0003] The present invention generally relates to deployable structuressuch as solar arrays and, more particularly, to solar array supportsystems.

[0004] 2. Description of the Related Art

[0005] In outer space applications, solar arrays are generally composedof a series of solar panels for generating electrical power for systemssuch as a spacecraft or the like. The conversion of solar energy intoelectrical energy through solar cells is an obvious choice for producingpower for such systems. The solar arrays are typically designed infoldable configurations in which the solar panels, supporting solarcells, are hingeably connected edge to edge lengthwise by variousattachment systems. Such solar arrays may also include reflectors toconcentrate the solar light upon the solar arrays. Reflectors areattached to the solar panels widthwise to the opposite ends of eachsolar panel.

[0006] The solar arrays are generally mounted on deployment yokes sothat they can be extended or retracted from the spacecraft. During thelaunch of the spacecraft, the solar arrays are put into a stowedconfiguration where the solar arrays are folded in zigzag fashionagainst the spacecraft. Once the spacecraft is in outer space, the solararrays are deployed into an extended configuration where the solarpanels and the reflectors are folded away into an operation position inwhich the solar cells face the sun.

[0007] Such solar arrays must be adequately designed to withstand theundesirable physical conditions of such space missions so that they canproperly function throughout their life time. Such undesirableconditions are generally mechanical and thermal stresses occurringduring the launching and during the operation of the solar arrays. Inthis respect, the solar panels supporting the solar cells must bedesigned to meet the certain thermal and mechanical stress and strainrequirements so as to protect the solar cells on them. As the solarcells are made of silicon or gallium arsenide materials, they arebrittle. In other words, the panels function as the mechanical and thethermal support of the solar arrays.

[0008] Currently, solar panels are constructed from aluminum honeycombsubstrates. In such structures, the honeycomb substrates are coveredwith carbon fiber face sheets on upper and lower surfaces of thehoneycomb substrates. The carbon fiber face sheets stiffen the honeycombsubstrate to increase the strength and the rigidity of the solar panelsin stowed or deployed configurations. However, such aluminum honeycombbase solar panels are heavy in the context of such space applications.Another drawback involves their deployment and stowing systems. Suchsolar arrays require complex deployment and stowing systems employingtension wires, springs, hinges and the like to facilitate the deploymentand stowing of the solar arrays.

[0009] As can be seen, there is a need for light weight, hightemperature resistant, stiff and resilient deployable structures such assolar arrays, reflectors, and thermal blankets.

SUMMARY OF THE INVENTION

[0010] The present invention provides a deployable structure, such as asolar array system, utilizing a carbon fiber mesh material as asubstrate. Due to its light-weight and flexibility, substratescontaining the carbon fiber mesh material can be applied, as an example,to deployable solar arrays and reflectors for air borne vehicles such assatellites for space based applications as well as to the solar arraysfor use in stratospheric platforms of an air borne vehicle such as anairplane.

[0011] In one aspect of the present invention, a solar array comprises amesh substrate and a plurality of solar cells disposed on a firstsurface of the mesh substrate. The mesh substrate is formed from amatrix of resilient fibers.

[0012] These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdrawings, description and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0013]FIG. 1 is a perspective view of a satellite using solar arraysaccording to an embodiment of the present invention;

[0014] FIGS. 2A-2B are schematic cross-sectional views of the solararrays shown in FIG. 1;

[0015]FIG. 3A is a schematic view showing the solar array of FIG. 2A ina fully retracted state;

[0016]FIG. 3B is a schematic view showing the solar array of FIG. 3A ina fully deployed state;

[0017]FIG. 4 is a magnified schematic illustration of the structure ofthe carbon fiber mesh substrate of this invention;

[0018]FIG. 5A is a schematic cross sectional view of a solar panelaccording to an embodiment of the present invention;

[0019]FIG. 5B is a schematic cross sectional view of a solar panelaccording to another embodiment of the present invention;

[0020]FIG. 5C is a schematic cross sectional view of the solar panelusing a carbon fiber mesh composite;

[0021]FIG. 6A is a cross sectional view of an embodiment of the solararray of the present invention;

[0022] FIGS. 6B-6D are schematic views illustrating deployed and stowedstates of the solar array shown in FIG. 6A;

[0023] FIGS. 7A-7B are schematic views illustrating stratosphericplatforms using an embodiment of the solar array of the presentinvention; and

[0024] FIGS. 8A-8C are schematic views illustrating solar reflectors ofthe present invention.

DETAILED DESCRIPTION OF THE INVENTION

[0025] Although the present invention is described in the context of asolar array system utilizing a carbon fiber mesh material as asubstrate, the scope of the invention is not so limited. Due to itslight weight and flexibility, substrates containing the carbon fibermesh material can be applied to deployable structures that supportobjects. What is meant by “deployable” is a structure that can bedeployed for use and stowed for non-use. Accordingly, these deployablestructures can include thermal blankets, solar arrays and reflectors forspace and land based applications. In particular, the present inventioncan be applied to the solar arrays for use in air borne vehicles such assatellites and stratospheric platforms of airplanes. In such uses, thecarbon fiber mesh material may be used in combination with othermaterials or laminates so as to enhance its applicability andversatility.

[0026] The carbon fiber mesh substrate of the present invention providesa material that is lighter and stiffer than the conventional substratesemploying aluminum honeycomb substrates or cores. Further, the carbonfiber mesh material of the present invention may withstand very highoperation temperatures, as high as 2500° C. This property significantlyalleviates the localized heating problems in solar arrays. Adding to theaforementioned advantages, the carbon fiber mesh material of the presentinvention is very resilient to bending. If the carbon fiber meshmaterial is folded into a compact shape, when released, it will springback to its original shape without leaving any fold line behind. Thepresent invention advantageously employs this property of the carbonfiber mesh to construct deployable solar arrays and solar reflectors.

[0027] Reference will now be made to the drawings wherein like numeralsrefer to like parts throughout. FIG. 1 exemplifies a spacecraft 100comprising a number of solar arrays 102 of the present invention in afully deployed configuration. The solar arrays may be mounted on aside-wall 104 of a body 106 of the spacecraft 100 using yokes 108. Inthis embodiment, although it is possible to use a multiplicity of thesolar arrays, for the purpose of clarity, only two of the solar arrays102 are shown on the spacecraft body 106. It is within the scope of thisinvention that the solar arrays 102 of the present invention may be usedwith any suitable structure including, but not limited to, astratospheric platform, satellite, space station or land basedapplications. The solar array 102 of the present invention may comprisea series of solar panels 110 comprising a plurality of solar cells 112on an upper side of the solar arrays 102. Solar panels 110 are separatedfrom one other at a folding section 114. As will be described below, thefolding sections 114 allow panels 110 to be folded relative to eachother in a zigzag fashion when the solar array is stowed. The solararray 102 of the present embodiment may also employ solar reflectors toincrease the power output of the solar cells. In this embodiment, a pairof solar reflectors 116 may be connected to each solar panel 110 usingrigid connectors 118 as in the manner shown in FIG. 1.

[0028] In FIG. 2A, in an exemplary cross-sectional view of the solararray 102, a number of solar panels 110 are shown separated by thefolding sections 114. The solar panels 110 may be comprised of a seriesof hybrid structures formed along a common and continuous substrate 120by sandwiching the substrate 120 between a series of discontinuous upperand lower support layers 122 and 124. In order to construct the solarpanels having such hybrid structure, a series of upper support layers122 or upper face sheets is disposed on a top surface 126 of thesubstrate 120 and between the folding sections 114. The solar cells 112may be placed on top of the upper face sheets 122. Similarly, a seriesof lower support layers 124 or lower face sheets are disposed on abottom surface 128 of the substrate 120 and between the folding sections114. At this point, the folding sections 114 can be further described asthe regions where the upper and lower face sheets 122 and 124discontinue and expose the underlying substrate 120. As will bedescribed more fully below, in this embodiment, the material of thesubstrate 120 is preferably less stiff than the hybrid structure of thesolar panels 110 which has the substrate 120 interposed between theupper and the lower face sheets 122 and 124.

[0029] Therefore, as illustrated in FIG. 2B, the solar array 102 can befolded along the folding section 114 when a bending force is applied,for example, in the direction of the arrow 130 and over one of thestiffened solar panels 110. However, due to the material properties ofthe substrate 120, when the bending force is removed, the solar arraysprings back to its straight configuration which is shown in FIG. 2A. Inother words, the bending force elastically deforms the folding sectionunder the deforming force, i.e., bending force in this case. Since thedeformation of the substrate material, which is exposed in the foldingsection 114, is elastic, i.e., temporary or non-plastic deformation,when the force is removed the folding section returns to its originalshape thereby deploying the solar array 102 without needing prior arthinge systems, springs and other deployment equipment. The dimensions ofthe substrate 120 depend on the application. Typically, the foldingsection 114 is approximately 3 to 4 times the thickness of thesubstrate.

[0030]FIGS. 3A and 3B show an implementation of the present invention.In FIG. 3A, in a stowed configuration, the solar panels 110 of the solararray 102 are folded against the body 106 of the spacecraft 100 (shownin FIG. 1) in a zigzag fashion by folding them along the foldingsections 114. In the stowed configuration, the solar panels 110 arefolded into a compact structure and secured (tied down) against thespacecraft until the time of the deployment. The stiff structure of thesolar panels 110 provides a rigid mounting base for the solar cells 112,which is less likely to deform and damage the solar cells. Once thespacecraft reaches the outer space orbit, as shown in FIG. 3B, the solararray 102 of the present invention is released and the solar panels 110are fully extended into a deployed configuration.

[0031] In the preferred embodiment, the substrate 120 comprises carbonfiber mesh material such as the one available from Energy ScienceLaboratory in San Diego, Calif. The carbon fiber mesh substrate materialof the present invention is a substantially light-weight material,although it is stiff, resistant to temperature extremes and resilient tobending. As illustrated in FIG. 4, the carbon fiber mesh material of thepresent invention is comprised of a network of randomly oriented carbonfibers 132 linked into a matrix 134. The fiber matrix may also have aplurality of voids formed between the carbon fibers 132. Accordingly,the substrate 120 material is distinguishable from the carbon fiber facesheets described above in the context of the prior art. Among otherthings, the substrate 120 material is not in a woven or unwoven matform. Further, the substrate 120 material is neither composed ofunidirectional or bidirectional fibers. As a result of this structure,the carbon fiber mesh material of the present invention may have a massof approximately 3 grams per square meter. Due to its carbon content(carbon is a refractory material), the carbon fiber mesh material cantolerate temperatures as high as 2500° C. This property of the carbonfiber mesh protects the solar arrays from the harmful effects oflocalized heating occurring during the operation of the solar arrays.

[0032] As previously mentioned, the carbon fiber mesh material can befolded under a bending force, but when the force is removed, thematerial recovers to its original shape without leaving a fold line. Inthis embodiment, the face sheets 122 and 124 may preferably be carbonfiber mats. Alternatively, Kevlar™ mats can also be used. The carbonfiber mats are generally reinforced carbon fiber composites impregnatedwith epoxy. The orientation of the carbon fibers in the mat iscontrolled so as to provide maximum possible tensile strength in a givendirection. As will be described below, multiple layers of face sheetscan be used with different carbon fiber orientations to derive stiffnessto bending along different axes.

[0033] In accordance with the principles of the present invention, thesolar panels 110 of the solar arrays may be fabricated in variousalternative ways. As illustrated in FIG. 5A, in order to form the abovementioned hybrid panel structure, the upper and lower carbon fiberface-sheets 122 and 124 are bonded over the top and bottom surfaces 126and 128 of the carbon mesh substrate 120 using a bonding material 136.The solar cells may be bonded to the upper face sheet 122 using thebonding material 136. In this embodiment a preferred bonding material isepoxy. As previously mentioned, the face sheets 122 and 124 stiffen thesubstrate in the hybrid structure of the solar panels 110.

[0034]FIG. 5B shows another embodiment of forming the hybrid structureof the solar panels using multiple face sheets. In this embodiment,after bonding the upper and lower face sheets 122 and 126 to the carbonfiber mesh substrate 120, face sheets 138 and 140 are bonded to the facesheets 122 and 124 respectively. The solar cells 112 are bonded to theface sheet 138 using the bonding material 136. As previously mentioned,the use of multiple layers of face sheets increases the stiffness of thesolar panels.

[0035] As shown in FIG. 5C, in another embodiment, a solar panel 142 maycomprise a laminated composite structure 144 or carbon fiber meshcomposite comprising a carbon fiber mesh substrate 146 interposedbetween the carbon fiber mats 148. Solar cells 150 are bonded to a topside of the carbon fiber mesh composite 144 using a bonding material 152such as those used in the previous embodiments. In this embodiment, itis within the scope of the present invention that the entire solar arraymay be formed from the carbon fiber mesh composite 144 without havingindividual solar panels 142.

[0036] As illustrated in FIGS. 6A-6D, for various solar arrayapplications, the stiffness of the carbon fiber mesh material can besufficient without the use of additional face sheet materials. One suchapplication would be an array composed of amorphous silicon cells. Sincethese cells are thinner and more flexible than conventional solar cells,the carbon fiber mesh alone can provide a sufficiently stiff substrate.As shown in FIG. 6A, a solar array 160 may comprise a carbon fiber meshsubstrate 162 and solar cells 164 bonded on top of the substrate 162using a bonding material 166. The bonding material 162 is the samebonding material used in the previous embodiments. FIGS. 6B and 6C showtwo possible stowing configurations for the solar array 160, a zigzagand a roll configuration respectively. As shown in FIG. 6D, whendeployed, the solar array 160 extends into its original shape withouthaving any fold or fatigue line.

[0037] As shown in FIGS. 7A and 7B, a stratospheric platform 170 maycomprise a number of solar arrays 172 to supply solar energy to anairplane 174. In this embodiment, the stratospheric platform 170 may becomprised of the wings of the airplane 174. As shown in FIG. 7B, thesolar array 172 comprising a solar array with the laminated compositestructure 144 described in FIG. 5C, or the solar array structure 160described in FIG. 6A can be applied over an upper skin 178 of the wing170. The lower mass and the stiffness properties of the carbon fibermesh material can allow overall reduction in the mass of the wing.

[0038] The solar reflectors 116 shown in FIG. 1 can be also fabricatedusing carbon fiber mesh material. As shown in FIG. 1, the solarreflectors are attached to solar panels 110 of the solar array 102 usingthe rigid connectors 118. Solar reflectors concentrate additional lightonto the solar cells to increase the power output of the solar array. Inthe prior art, the reflectors are held in tension by a system of springsand rods and attached to the solar panels using hinge systems. Thedeployment of such prior art systems is difficult to implement.

[0039] As illustrated in FIG. 8A, in one embodiment, the reflectors 116may be comprised of a carbon fiber mesh substrate 180 having areflective layer 182 on top of the substrate 180. The reflective layermay comprise vapor deposited aluminum (VDA) Kapton™ material. As shownin FIG. 8B, in a fully deployed configuration, the reflectors 116 areattached to the solar panels 110 using rigid connectors 118, therebyeliminating prior art hinges and tension wires. As shown in FIG. 8C, thesolar reflectors 116 may be stowed by bending the reflectors toward thetop and the bottom of the solar array 102. The tendency of the carbonfiber mesh to spring back to its original shape makes it possible torigidly mount the reflector to the solar array and stow the reflector bybending the carbon fiber mesh reflectors.

[0040] It should be understood, of course, that the foregoing relates topreferred embodiments of the invention and that modifications may bemade without departing from the spirit and scope of the invention as setforth in the following claims.

We claim:
 1. A solar array comprising: a mesh substrate, the meshsubstrate being formed from a matrix of resilient fibers; and aplurality of solar cells disposed on a first surface of the meshsubstrate.
 2. A solar array comprising: a mesh substrate; said meshsubstrate consisting essentially of a matrix of randomly orientedresilient fibers and having a first surface; and a plurality of solarcells disposed on the first surface of the mesh substrate.
 3. The solararray of claim 2, wherein said matrix further comprises a plurality ofvoids between said fibers.
 4. The solar array of claim 2, wherein saidmesh substrate is characterized by a mass of approximately 3 grams persquare meter.
 5. A solar array comprising: a mesh substrate; said meshsubstrate consisting essentially of a matrix of carbon fibers; saidcarbon fibers being randomly oriented in said matrix; and a plurality ofsolar cells disposed on said mesh substrate.
 6. The solar array of claim5, wherein said matrix further comprises a plurality of voids betweensaid fibers.
 7. The solar array of claim 5, wherein said mesh substrateis characterized by a mass of approximately 3 grams per square meter. 8.The solar array of claim 5, further comprising a support layerinterposed between said mesh substrate and said plurality of solarcells.
 9. A solar array comprising: a mesh substrate; said meshsubstrate consisting essentially of a matrix of resilient carbon fibers;said matrix being characterized by a plurality of voids; said carbonfibers being randomly oriented in said matrix; and a plurality of solarcells disposed on said mesh substrate.